1. Technical Field
This invention relates to gas turbine engines in general, and to methods and apparatus for cooling a rotor blade or stator vane in particular.
2. Background Information
Efficiency is a primary concern in the design of any gas turbine engine. Historically, one of the principle techniques for increasing efficiency has been to increase the gas path temperatures within the engine. The increased temperatures have been accommodated by using internally cooled components made from high temperature capacity alloys. Turbine stator vanes and blades, for example, are typically cooled using compressor air worked to a higher pressure, but still at a lower temperature than that of the core gas flow passing by blade or vane. The higher pressure provides the energy necessary to push the air through the component. A significant percentage of the work imparted to the air bled from the compressor, however, is lost during the cooling process. The lost work does not add to the thrust of the engine and therefore negatively effects the overall efficiency of the engine. A person of skill in the art will recognize, therefore, that there is a tension between the efficiency gained from higher core gas path temperatures and the concomitant need to cool turbine components and the efficiency lost from bleeding air to perform that cooling.
There is, accordingly, great value in maximizing the cooling effectiveness of whatever cooling air is used. Prior art coolable airfoils typically include a plurality of internal cavities, which are supplied with cooling air. The cooling air passes through the wall of the airfoil (or the platform) and transfers thermal energy away from the airfoil in the process. The manner in which the cooling air passes through the airfoil wall is critical to the efficiency of the process. In some instances, cooling air is passed through straight or diffused cooling apertures to convectively cool the wall and establish an external film of cooling air. A minimal pressure drop is typically required across these type cooling apertures to minimize the amount of cooling air that is immediately lost to the free-stream hot core gas passing by the airfoil. The minimal pressure drop is usually produced through a plurality of cavities within the airfoil connected by a plurality of metering holes. Too small a pressure drop across the airfoil wall can result in undesirable hot core gas in-flow. In all cases, the minimal dwell time in the cooling aperture as well as the size of the cooling aperture make this type of convective cooling relatively inefficient.
Some airfoils convectively cool by passing cooling air through passages disposed within a wall or platform. Typically, those passages extend a significant distance within the wall or platform. There are several potential problems with this type of cooling scheme. First, the heat transfer rate between the passage walls and the cooling air decreases markedly as a function of distance traveled within the passage. As a result, cooling air flow adequately cooling the beginning of the passage may not adequately cool the end of the passage. If the cooling air flow is increased to provide adequate cooling at the end of the passage, the beginning of the passage may be excessively cooled, consequently wasting cooling air. Second, the thermal profile of an airfoil is typically non-uniform and will contain regions exposed to a greater or lesser thermal load. The prior art internal cooling passages extending a significant distance within an airfoil wall or a platform typically span one or more regions having disparate thermal loads. Similar to the situation described above, providing a cooling flow adequate to cool the region with the greatest thermal load can result in other regions along the passage being excessively cooled.
What is needed, therefore, is a method and apparatus for cooling a substrate within gas turbine engine that adequately cools the substrate using a minimal amount of cooling air and one that provides heat transfer where it is needed.